Helicopter rotor construction



March 18, 1952 J. A. J. BENNETT HELICOPTER ROTOR CONSTRUCTION 2 SHEETS-SHEET l Filed Aug. 5, 1946 March 18, 1952 J. A. J. BENNETT HELICOPTER ROTOR CONSTRUCTION Filed Aug. 5, 1946 2` SHEETS- SHEET 2 Patented Mar. 18, 1952 UNITED STATES PATENT OFFICE HELICOPTER Ro'ron CONSTRUCTION James Allan Jamieson Bennett, Dumbarton, Scotland Application August 5, 1946, Serial No. 688,451 In Great Britain January 11, 1946 path plane, therefore, being inclinable in flight with respect to the plane normal to the axis of the rotor hub.

An object of the invention is to provide an arrangement for eliminating or minimizing the unstable oscillation of the aircraft, especially in hovering night, when it is displaced angularly, in pitch or in roll, by a gust. More particularly, an object of the invention is to improve the dynamic stability of a helicopter, when hovering, by displacing automatically the centre of pressure of the rotor thrust whenever the tip-path plane is displaced angularly with respect to the plane normal to the axis of the rotor hub, the moment of the rotor thrust resulting from the displacement of the centre of pressure being a stabilizing one, i. e. in the opposite sense to the moment of the rotor thrust resulting from the angular displacement of the tip path plane and, therefore, of the rotor thrust which isnormal to this plane.

In a rotary wing aircraft according lto the invention, particularly a helicopter, each blade is mounted on a flapping pivot located on the side of the rotor hub axis remote from the blade. An inclination of the tip-path plane with respect to the plane normal to the hub axis thereby displaces automatically the centre of pressure of the rotor thrust in the direction required for improving the dynamic stability of the helicopter.

The invention will be explained with reference to the accompanying drawing in which:

Figure 1 is a diagrammatic side elevation and Figure 2 is a diagrammatic plan.

Figure 3 is a sectional side elevation of one embodiment of the invention, taken on the line III-III, Figure 4, parts being omitted for the sake of clearness.

Figure 4 is a plan of the embodiment shown in Figure 3, parts being shown in section on the line IV--IV, Figure 3, and

Figure 5 is a detail section on theline V-V, Figure 4.

Figure 6 is a diagrammatic side elevational view showing the relationship of the factors involved in the invention.

Referring now to Figures 1 and 2 of the drawings, I indicates the rotor head of a rotary-wing aircraft, 2 indicates a drag link, 3 a blade pivoted to the drag link 2 by means of a drag hinge 4, While 5 indicates the flapping hinge by which the drag link 2 is pivoted to the rotor head I.

2 If the metacentre be dened as the point M, on the hub axis 6, through which the rotor thrust always passes when the tip-path plane 'l 'is inclined slightly as at 'l' to the plane 'l normal to the hub axis 6, the location of the blade napping pivots 5 in accordance with the present invention ensures that the metacentre M is below the plane containing the flapping pivots 5. Hitherto, rotary wing aircraft having rotor blades mounted on flapping pivots intersecting the hub axis, or located on the same side of the hub axis as the blade, have resulted in a metacentre in or above the plane containing the flapping pivots.

If the flapping pivot axis 5 be located at a distance from the rotor hub axis B on the side of this axis remote from the blade 3, and if the blades 3 have a coning angle C, the metacentre M is approximately at a distance f cot C' below the plane 8 in which the flapping pivots 5 lie, i. e. below the position Mo of the metacentre when f=0. More particularly the metacentre can be defined as the point on the rotor axis located at a -distance ,f cot C' below the plane of the blade flapping pivots, Where f is the distance of each flapping pivot from the rotor axis and C is the blade coning angle in normal hovering.

C is usually a small angle, being approximately the angle of which the tangent is the ratio of lift to centrifugal force, or the value of the coning angle which prevails during normal hovering. The distance MMO (viz. f cot C) will usually amount to several feet if f be a few inches.

Hence, if G be the nearest point on the hub axis 6 to the centre of gravity C. Gr. of the aircraft, the metacentre height GM may be decreased from GMO to Zero by increasing f from zero to the value (GMO) tan C, in which case the static stability of the air craft would bereduced substantially to zero. This would be undesirable and therefore the distance f is preferably chosen so that f cot C is less than GMO. In other words, the improvement in dynamic stability obtained by increasing f is accompanied by a reduction in static stability and there are therefore lower and upper limits within which f should preferably be chosen. It can be shown mathematically that these limits are given by the inequality where h=GMo and lc is the radius of gyration of the aircraft about the axis of pitch or roll.

Considering a rotary wing aircraft as shown in Fig. 6, axes Or; Oy, Oz are taken forward, side- Ways, and downwards, respectively, through the centre of gravity. The aircraft is considered to be hovering in equilibrium, with the rotor providing a thrust equal to the weight Mg of the aircraft. Then suppose that the aircraft is subjected to a small disturbance,v causing the aircraftito. acquire velocities :z: and .e horizontally and vertically, respectively, and an angular displacement in pitch of 0.

For equilibrium of the forces atv O inytheV Om and Oz directions and of longitudinal moments about the Oy axis, the equations governing the disturbance are:

B=Hz1Tl1 where Trotor' thrust Y Illbackward force on rotor-'normal to T M=`mass:of aircraft* B=n1oment of inertia of the aircraft `about they axis. Z1=distanceqfrom centre of gravity to the rotor r headf =a00 and o ao l a is here' treated, as a small quantity, and, as T is of.' theA orderof'Mg, and H is a small quantity of the first order proportional to the disturbance velocity of' the rotor-head (a'c- 116), the term H0 has been neglected, and only first order terms retained.

If the oscillation of the tip-path plane is reduced in amplitude by an times the oscillation of the body, an being less than 1, the angle denotes the reduction in angle of oscillation of the tippath plane when the amplitude of oscillation of the body is denoted by the Greek letter 0.

Since the. longitudinal force'on a rotor is proportional to the velocity, We may write where R=tip speed and R=rotor radius.

The-soluti'onbf these two. simultaneous. equationsrr is. obtained by assuming ac-A1ew and =A2ewb giving, two homogeneous. linear equationsin A1A and A2,;as follows:

MwA1+ (MJFMQAF o ne R Eliminating A1 and A2, we obtain the frequency equation for w:

where w is the frequency of the oscillation.

The condition for stability is that the roots of the frequency equation should be real and negative, or imaginary with their real parts negative, and Routh has shown in his textbook Advanced Rigidv Dynamics (1892) that this condition is satisfied provided the coefcients A, B, C and D are positive and, in addition, BC AD.

Where k1 is the radius of gyration of the aircraft.

Therefore.

Similarly the samek inequality can be shown to apply for stability about the axis of roll.

In the embodiment of the invention illustrated by Figures 3, 4 and 5 the rotor headv I is mounted in ball bearings 9 in a fixed hub II), the upper part of the rotor head I being substantially triangular in plan, as shown in Figure 4, and carrying pairs of lugs II, II between which drag links 2 are pivoted on vthe flapping hinge axes 5, The blades 3 are pivoted. to the opposite ends of the drag links 2 on drag hinges 4' and each flapping hinge 5 is situated on that side of the rotor hub axis 6 remote from the corresponding blade 3.

In Figure 4 the positive torque position of a blade 3 is indicated at I 2 and` its zero torque position at I3.

I claim:

A helicopter comprisingV a bladed lift rotor wherein each blade is independently pivoted on a flapping pivot located on that side of the rotor hubraxis remote from the major portion of the corresponding blade, eachof said pivots having only a single pivot axis, and wherein the distance'of each pivot axis'fromthe rotor hub axis lieswithin the limits'expr'essed by the inequality REFERENCES CITED- The following references are of record in the le of this patent:

UNITED STATES PATENTS Name Date Rothenhoefer Oct. l1, 1938 Number 

